The invention relates to air sseals for use in turbo-machinery, and, more particularly to air seals with increased cooling and flexibility.
In turbo-machinery, such as gas turbine engines for aircraft, a flow of high pressure gas is directed onto a plurality of turbine blades mounted upon rotatable disks. The gas imparts momentum to the turbine blades in a manner which transforms the kinetic energy of the gas flow into torque to operate rotating elements. Gas turbine efficiency depends to a great degree upon directing the high pressure gas (working fluid) onto the plurality of turbine blades while restricting the high pressure gas from bypassing around the tips of the turbine blades. When gas is allowed to bypass the blades, the gas turbine engine looses efficiency due to the unapplied loss of useable kinetic energy.
In view of the above, gas turbine engines incorporate a turbine casing having air seals which surround the turbine blades and define the outer flow path of the pressurized gas in the vicinity of the blades. Clearance gaps between the radial blade tips and the air seals permits a portion of the working fluid to bypass the rotating blades.
In order to minimize efficiency losses due to unrestricted flow of the working fluid it is important to minimize the air gap between the rotating turbine blades and the turbine casing. If the air gap is made too small, however, another problem arises. During engine startup or acceleration turbine blade temperature increases rapidly, which in turn results in an increase in turbine blade size due to the thermal expansion. The turbine casing and the other non-rotating elements of the turbine such as the air seals, do not heat as rapidly as the turbine blades and therefore do not expand as quickly as the turbine blades. This can result in destructive contact between the turbine blades and the casing. This problem is particularly important in turbine machinery where the critical clearance required to maintain high efficiency is quite small. As a result of this problem, modern turbine engines make use of annular air seal shrouds supported adjacent to the rotating turbine blades by the turbine case. The shrouds' internal surface is coated with a ceramic material that withstands high turbine temperatures and permits some non-destructive rubbing between the turbine blades and the shroud in order to help maintain very small clearances.
It has been an object in modern gas turbine engines to provide air seals that can withstand thermal and phyical stresses in close proximity to both a very high temperature, high pressure gas stream and the rotating high temperature turbine blades. This has been efficiently done through the use of cooling air being directed onto the outer, or back surface, of the air seal itself. This air is generally relatively cool compressor air that has been bypassed around the engine combustor and directed to the air seals. This air helps control air seal temperature (and thermal expansion) by heating the air seal during engine startup and cooling it during steady state operation.
The bypass air cooled seal therefore allows the turbine to operate at elevated temperatures by providing a cooling system which prevents damage to the air seal due to the effects of high turbine temperatures.
This system has generally worked well for the current generation of aircraft engines. Certain shortcomings, however, have arisen in recent years due to the drive for more efficient and more powerful aircraft engines. In order to raise the efficiency of a gas turbine engine it is generally necessary to raise the turbine operating temperature. It is therefore not unusual for modern turbine operating temperatures to be in the range of 2000.degree. F. at high power settings. These higher temperatures have resulted in seal buckling and deterioration that has decreased air seal life and resulted in lowered long term engine performance due to increased air gaps between turbine blades and the air seals (shrouds).
An example of conventional turbine air seal is disclosed in U.S. Pat. No. 3,583,824 to Smuland et al. The Smuland patent discloses an air seal wherein bypass air is directed through a perforated baffle 54 and impinges on the back of a turbine shroud (air seal) 26. The air therefore cools the center section of the shroud and exits through a hole 66 to eventually join with the air stream passing by turbine blades 24. This type of cooling system has been generally successful.
New problems, hwoever, have recently developed in high temperature engines. The Smuland device makes no provision for direct cooling of either axial end of the shroud 26; the cooling air goes mainly in the center two-thirds of the shroud. Turbines utilizing seals such as these have had problems with shroud edges becoming burnt or buckled. To solve these problems, angled holes have been drilled through the shroud material to deliver cooling air to the edges of the air seal. This in turn has resulted in an increased cooling air flow that lowers engine efficiency. In addition, manufacturing air seals with these small angled holes is quite expensive and difficult. Further, the holes are subject to blockage since they are quite long and relatively small diameter. When the holes are blocked shroud failures can occur due to overheating.
Secondary air seals have also been added to this type of design to control bypass air flow and minimize its effect on efficiency. Thus it can be seen that turbine air seals can be quite complex and difficult to manufacture.
A primary object of this invention therefore, is to provide a cooling system for turbine air seals that will allow turbine operation at increased temperatures without damage to the seal material.
Another object of this invention is to provide a cooling system which provides positive seal edge cooling.
A further object of this invention is to increase air seal flexibility to allow for seal thermal growth and contraction without buckling.
Yet another object of this invention is to eliminate difficult machining operations for air seal manufacture.